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Liquid propulsion ppt

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Liquid propellant rocket enginesGeneral characteristicsLiquid propellant rocket engines are mostly widely used rocket engines because of manyadvantages that liquid propellants have. The first rockets used solid propellants because of thesimplicity of their construction (just a barrel with gunpowder), but such engines were difficultto control. Chemistry and physics of combustion were undeveloped, combustion wasunpredictable and it was nearly impossible to control it. Liquid rocket engines (LRE) werevery promising: their thrust could be controlled by dosing propellant flow ratio with valves.Although nowadays the techniques of solid propellants have enormously advanced, liquidpropellants retain their importance for the rocketry.The main advantages of liquid propellants:• high specific impulse;• high thrust;• high thrust to weight ratio of the rockets;• easy to control.However, they also have certain disadvantages:• complexity of the construction;• it is impossible to achieve very high specific impulse;• they are difficult to scale, complexity grows quickly with growing the thrust;• some propellants are highly toxic (hydrazine and its variants) or cryogenic (hydrogen). 2

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The working principle of all liquid rocket engines is transformation of the potentialchemical energy of liquid propellants to kinetic energy of the exhaust gases. It is important tomention that LRE is only a part of the propulsion system; other parts are tanks, plumbing,hydraulics, framework etc.Types of propellantsThere are two basic types of LRE propellants: monopropellants and bipropellants. Monopropellants are liquids which may be stored in a single tank (and remain stable).They are decomposed releasing energy in presence of a catalyst. Among suchmonopropellants are hydrogen peroxide H2O2 and hydrazine N2H4. Hydrogen peroxide isused, for example, as propellant for the turbopumps in RD-107/108 engines in the first andthe second stages of the Soyuz launch vehicles. It was also used in the Mercury mannedspacecraft, in the Centaur upper stage (in the ullage and attitude control motors) etc.However, this propellant slowly decomposes by itself, so it cannot be hold for years and thuscannot be used in spacecraft with long lifetimes. Hydrazine is widely used in maneuveringthrusters or main engines of spacecraft, also in descent engines. For example, hydrazine wasused in the thrusters of the Voyager spacecraft, in the descent engines of the Viking and thePhoenix Martian landing probes etc. Hydrazine is stable and may be hold for years. The main advantages of monopropellants is that they need only one tank and that theignition system is not needed, they react by themselves in presence of catalyst. Due to lowtemperatures in the thrust chamber (combustion chamber) they may work for a long time(for hours) and be restarted (may give thousands of very short pulses). That makes them idealfor thrusters. Their thrust may vary from tens of grams to several kilograms. The maindisadvantage is their low specific impulse (Isp), for hydrazine mostly Isp < 250 s. 3

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Bipropellants consist of fuel and oxidizer that should be hold separately and react in thethrust chamber when ignited. Bipropellant mixtures divide in practice by their properties into hypergolic and cryogenicpropellants (although the first does not exclude the second). Hypergolic mixtures ignitespontaneously when brought in contact. Thus, no complex ignition system and startingprocedure is needed (thus, multiple restarts are simply feasible). The process of combustion ofhypergolic propellants is also more stable, so engines are more simple to develop and lesslikely to be destroyed at work. Hard starts are less likely with hypergolic propellants (hardstart occurs when the ignition in the thrust chamber takes place at presence of excessiveconcentration of propellants, thus instantaneous overpressure is established and it may lead toexplosive destruction of the engine). The most widely used hypergolic mixture is hydrazine or its variants(monomethylhydrazine /MMH/, unsymmetrical dimethilhydrazine /UDMH/, aerozine /50% hydrazine + 50% UDMH/) as fuel and dinitrogen tetroxide (N2O4) (or nitric axid/NO/ in earlier applications) as oxidizer. The great advantage of this propellant is that the fueland the oxidizer both are liquids at the normal conditions, they are not cryogenic, so they maybe stored at the common temperatures. This makes them ideal for military applications(ICBM may stay for arbitrary time in its silo or stored) as well as for spacecraft with a longlifetime (like interplanetary probes). This is the reason why they have been used on manyICBMs which later have been converted into launch vehicles: the Titan, the Strela and theRokot (ICBM UR-100/100H), the Dnepr (R-36M). These mixtures have quite high Isp(vacuum values 300 – 320 s and even more), so they are most common on spacecraft,although need more complex engines when that based on the hydrazine as monopropellant.The Cassini spacecraft uses hydrazine monopropellant thrusters for small attitude maneuvers,but burns 4

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N2O4/MMH in its main engine. Orbital engines of the Space Shuttle use N2O4/MMH, ApolloLM and CSM used N2O4/aerozine, Luna E8 series (Lunokhod, soil sample missions) usedN2O4/UDMH. The highest disadvantage of this mixture is that it is highly toxic (both fuel andoxidizer), so it should be handled very carefully. The flame of hypergol is nearly colorless,slightly blue. Plumes of even powerful engines burning this propellant is nearly invisible. Cryogenic propellants include mixtures at least one component of which (fuel or oxidizeror both) need low temperatures to liquefy. One of the most common such mixtures isLOX/kerosene (some special technological processes are used to produce kerosene for rocketindustry, the resulting fuels have different names; in US RP-1 is a popular type of kerosenefuels, in USSR sintin was used). LOX/UDMH is also an option. LOX, one of the mostwidespread oxidizers, boils at –1830C, thus being moderately cryogenic: this temperature ishigher than the boiling point of air (– 1940C), which is produced in industrial quantities. So,air does not liquefy on cold walls of LOX tanks, and extensive termostating is not needed. ForLOX/kerosene propellants, in vacuum Isp ≈ 340 s, and it is very widespread propellant forlaunch vehicles, specially for the first stages. It is used on all stages of the Soyuz and theZenit, on the first stage of the Atlas V, it was used on the first stages of the Saturn I/IB, theSaturn V and the Energia. Due to its better energetic characteristics than that of thehypergols, LOX/kerosene is also used on the Blok DM, the 4th stage of the Proton (3-stageN2O4/UDMH heavy rocket). Cryogenic propellants cannot be held for a long time, and thetanks should be insulated. In space, some amount of the propellant boils out, and if some timepasses between multiple burns, the loss may be significant, so instulation is needed. Plumes ofLOX/kerosene engines contain many carbon particles and thus are bright yellowish. The most energetic propellant that is used in practice is LOX/LH2, it may have Isp > 450 s(vacuum value). Liquid hydrogen is highly cryogenic fuel, having the boiling point at –253 0C. 5

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So it is difficult to use, since continuous thermostating is unavoidable. Tanks containing LH2should have extensive thermal protection (which leads to increase of their weight), the fuelcannot be hold in space for a long time since it boils out. Extremely low temperatures of LH2lead to changes of physical properties of metals which contact with this fuel, metals saturatewith H2; these factors should be taken into account at building of a LOX/LH2 rocket engine,plumbing and tanks. Because of very low density of LOX/LH2 (~280 kg/m3), it needs largetanks, which also increase the mass of stages with this propellant. In addition, Isp of suchengines significantly drops in the atmosphere. So, LOX/LH2 is used primarily on the upperstages of the launch vehicles. The first stage where this propellant was applied was theCentaur upper stage for the Atlas launch vehicle. Later it appeared on the upper stage S-IV(Saturn I) and S-IVB (Saturn IB and Saturn V), as well as on the second stage S-II of theSaturn V. The Space Shuttle has become the first spacecraft where LOX/LH2 is used on thestage working from the sea level. Later it appeared on the second stage (also working fromthe sea level) of the Energia. Today it is a common propellant on upper stages of launchvehicles: Delta, Atlas, Ariane, GSLV (India), CZ-3 (China). It is also used on the first stagestarted on the sea level, but is aided by strap-on solid rocket boosters: Ariane, H-II (Japan).There is only one full-cryogenic launch vehicle which burns this propellant on all stages, thatis the Delta IV. The plumes of LOX/LH2 engines are nearly invisible.Engine designsTo introduce propellants into the thrust chamber, two principle designs of LRE are used:pressure-fed and pump-fed. The first one is the most simple and reliable, the second oneenables to get higher specific impulse. 6

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Pressure-fed enginesIn a pressure-fed LRE, the propellants are forced to the thrust chamber by pressure of gaswhich pressurizes the tanks. For pressurization, a separate gas supply is provided. So, there isa special tank with pressurizing gas onboard (helium is commonly used for this purpose). The greatest advantage of pressure-fed engines is simplicity and thus reliability of thisdesign: contrary to pump-fed engines, no complex turbopumps are needed, no gasgenerators etc. Such engines contain much less parts and much less moving parts, so thereare much less things that might fail. The procedure of engine cut-off and restart is also verysimple: there is no need to stop and restart the turbopumps, its enough to close or open thevalves, and the propellant flow to the thrust chamber ceases or recommence (if the propellantis hypergolic, even no ignitor is needed). To avoid the pressurizing gas to cool down due toexpansion inside the tanks, it is often warmed up in the heat exchanger. The advantages of this solution make pressure-fed engines ideal for applications wherereliability and simplicity are important, as well as capability for multiple restarts. This is thereason why all engines of the Apollo CSM, as well as all engines of the Apollo LM werepressure-fed. Shuttle orbital maneuvering and control engines are pressure-fed as well.Maneuvering and attitude control thrusters of satellites and space probes are mostly pressure-fed since they are restarted thousands of times. However, this design have two principle disadvantages (mutually related). Specificcharacteristics of rocket engines depend on the pressure in the thrust chamber. But the 7

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pressure in the thrust chamber cannot exceed the pressure in the the tanks (actually, thepressure in the tanks should be higher). To withstang high pressure, large tanks should havemore robust and heavier construction. The larger is the tank, thicker should be its walls tobear the same pressure. Thus, pressure-fed systems are generally limited by chamberpressures of ~10 bar (on the Apollo SM it was 7 bar, on the LM Ascent Stage 8.4 bar). Theyare rarely applied on first stages due to large size of the tanks (however, the engine on the 2 ndstage of the Delta II rocket is pressure-fed). To avoid additional pressure in the tanks,regenerative cooling jacket is often avoided, that obliges to use ablative and radiativecooling.Pump-fed enginesPump-fed systems do not have the limitations of the pressure-fed systems. In this design, thepropellant is forced into the thrust chambers with dedicated pumps. The required efficiencymay be provided only with centrifugal pumps, herewith the pump should rotate at tens ofthousands rpm. Only a turbine is capable to ensure such speeds, so the natural solution is aturbopump. A turbopump consists of one or more pumps often mounted on the same shaftwith a driving turbine. The turbine is driven by gas flow, the gas may be produced in a gasgenerator by preburning some amount of the propellant, by burning a separate propellant(like hydrogen peroxide in the RD-107/108 engines on the Soyuz) or by gasification of somepropellant in the cooling jacket of the thrust chamber and the nozzle. The pumps may bemultistage. The turbopump assembly may include also booster pumps, which are added tounload principle pumps and to increase the pressure in the gas chamber (these pumps may bedriven by a hydraulic turbine powered by liquid from a high-pressure line, but also by themain turbine). Turbopump assembly is the most complex part of the engine, since it should 8

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have enormous productivity and work in harsh conditions (the turbine is driven by very hotgases and rotates very quickly). For example, the turbopump of RD-170/171 (the mostpowerful LRE ever produced, Energia & Zenit launch vehicles, LOX/kerosene) has a massflow rate of ~2.4 tons/s, it provides the pressure in the thrust chamber of ~250 bar, the powerof the turbine is ~200 MW, it rotates at ~14 000 rpm. The pressure of gases driving the turbineis ~500 bar, their temperature is ~5000C. At the same time the turbopump should be compactand lightweight (the mass of whole RD-170 is about 10 tons). So high characteristics arepossible only because the lifetime of such assemblies is only tens or hundreds of seconds.However, there exist pump-fed engines of multiple use which may work for hours, berestarted and continue to work after revision. An example of such engines is the SpaceShuttle Main Engine (SSME). The obvious advantage of pump-fed engines is that they may provide very high pressuresinside the thrust chamber and so their specific impulse is high. In spite of their complexity,they may be compact enough and be lighter than pressure-fed engines with their pressurizinggas vessels and thick propellant tanks; thanks to their efficiency, they make it possible tospend less propellant. Their complexity is the highest disadvantage, since complex turbopumpassemblies tend to be more expensive and less reliable than pressure-fed designs. However, ifefficiency is critical, pump-fed design is a natural solution. Turbopumps are used on all stagesof launch vehicles, but also on spacecraft. The Soviet lunar probes E8 (Lunokhods, soilsample missions) used pump-fed design, and the Soviet lunar module for the mannedexpeditions as well (since weight was critical). The space stations Salyut, the Soyuz mannedspacecraft have been provided with pump-fed engines. 9

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There are several designs of pump-fed engines. The most spread is the gas generatorcycle. In these engines the turbine of the turbopump is powered by gas resulting from burningsome of propellant in the gas generator (also called preburner sometimes) – a special smallcombustion chamber. In some designs there may be two gas generators (like the RD-170/171), sometimes each gas generator provides gas for separate turbines (of fuel andoxidizer). The mixture in the gas generator is ordinarily very fuel-rich or oxidizer-rich inorder to keep the temperature reasonably low and not to damage the blades of the turbine(actually, only small amount of the propellant burns, the rest is only gasified). After theturbine, the gas is ejected, either through the main nozzle either through a special nozzle. Dueto its low temperature, its contribution to the engine thrust is quite low, so it is nearly“wasted” for the thrust. Several percent of the propellant are lost. However, sometimes thisgas is used in steering nozzles or may participate in film cooling of the main nozzle (like inthe F-1 engine of the Saturn V). To improve efficiency of the engine, another version of this cycle is used, that is the so-called staged combustion cycle (or closed cycle). The main difference of this cycle is thatthe gas after turbine is not dumped, but is returned to the thrust chamber. So, all propellantand all heat pass through the thrust chamber and nothing is wasted. The disadvantage of thissolution is that the turbine have to do work against the pressure of the gases which it shouldpress into the thrust chamber. So the efficiency of the turbine drops, and it needs more powerto work. Thus, it works in worse and more harsh conditions, the plumbing of hot gases ductsis much more complex, as well as the control. So such engines are generally more complex,more expensive and less reliable. They are very sensitive to productional quality and toexternal particles that may occasionally get into the ducts, turbines and pumps. But the gainof Isp may be so high that this design makes sense. It first appeared in the USSR, and theyhave a long tradition of building engines of the closed cycle. For example, the RD-170/171applies the 10

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closed cycle (contrary to the F-1), as well as the SSME of the Shuttle. In most cases only small amount of the propellant is gasified in the gas generator (and inthe gas generator cycle it cannot be else to avoid excessive loss of propellant). But in somedevelopments the full amount of the fuel and the oxidizer passes through the turbine (the so-called full flow staged combustion cycle). It enables to reduce the temperature of the gas andthe rotation velocity of the turbine, since the it is driven by larger mass. The lifetime andreliability grow. Of course, two separate gas generators and turbines are needed for the fueland the oxidizer. However, separate systems for both components are usual for LOX/LH2engines, since the components have very different physical properties (density on the firstplace), so it is difficult to provide optimal characteristics for them in a single assembly. Sometimes it is possible to get rid of the gas generator assembly at all (gas generator is asmall combustion chamber by itself, with its own nozzle ejecting gas into the turbine, so it is acomplex unit). This is the expander cycle design. In this cycle the gas for driving the turbineis produced from the fuel vaporized in the cooling jacket of the thrust chamber and the nozzle.A gas generator is sometimes used to start the engine. This cycle may be opened or closed. Inthe opened cycle, only a small portion of the fuel is used to drive the turbine and thereafter itis dumped. In the closed cycle the fuel is redirected into the thrust chamber after leaving theturbine. Although the close cycle saves fuel, the open cycle enables higher pressure drop onthe turbine which increases its efficiency and enables to raise the pressure in the chamber.This leads to higher Isp (this is the case of LE-5A/B on the second stage of the Japanese H-IIrocket, LE-5 used the gas generator cycle). The famous RL-10 and its modifications on theCentaur upper stage use the expander cycle. Generally, the expander cycle is mostly appliedin LOX/LH2 engine since fuel is ordinarily used for regenerative cooling (oxidizer is tooreactive) and LH2 has low boiling point and is very effective as reaction mass. 11

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Thrust chamber and nozzleThe thrust chamber is the principle component of the rocket engine, the propellant is injectedinto it and burns, transforming into hot gases that escape through the nozzle (the bell). Thethrust chamber assembly consists of the following main components: the thrust chamberbody, the nozzle (the chamber narrows to its end, the narrowest part of it is called throat, andbehind the throat it expands again, forming the nozzle; the end part of the nozzle is calledextension), the injector. In the case of regenerative cooling, the body and the nozzle may becombined with a cooling jacket.Cooling of the thrust chamber and the nozzleSince gases in the thrust chamber have very high temperatures, its walls should be cooled, aswell as the walls of the nozzle. Without cooling, the walls cannot withstand such temperaturesfor a long time. There are two principles of cooling: passive and active cooling. Differentmethods may be applied simultaneously in different parts of the chamber and the nozzle. Passive cooling includes ablative and radiative methods. Ablative cooling means that thewalls are covered with substance called ablation, which have high heat capacity and absorbsheat by transforming itself chemically and/or physically. The ablation burns slowly andremoves heat with the gases created in this process. This method is limited by timespan andby temperature: ablative materials cannot withstand very high temperatures and since they aregradually removed, ablation works only for a limited time. However, the highest advantage ofthis method is simplicity and reliability, so it is sometimes applied even on large engines, likeRS-68, the most powerful LOX/LH2 engine in use (Delta IV). Radiative cooling is the process when the hot wall loses heat by radiation. Being very 12

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simple, this method is limited to thin surfaces with relatively moderate incident heat fluxes. Ifthe heat flux is very intensive, the equilibrium temperature of such wall with only radiativecooling is too high and the wall may be damaged. If the wall is thick, its hot side is damagedbefore the heat diffuses to the cold side. So, radiative cooling is mostly applied to cool nozzleextensions, as well as in small maneuvering engines, where the heat of short burns is absorbedby a massive conductive wall of the chamber (made from cooper alloy, for instance) and isirradiated between the burns. Active cooling methods include regenerative cooling and film cooling. Regenerativecooling means that a flow of cold propellant is organized along the the hot wall and thepropellant carries away excessive heat. Thus, the walls of the thrust chamber and the nozzlehave a cooling jacket with a propellant flow inside (fuel is commonly used). The propellantis directed into the jacket from the tank before it is injected into the thrust chamber. There aredifferent ways to build the cooling jacket. The simplest way is two walls, inner and outer,separated by a folded metal sheet, with propellant flowing along the folds. This design havebeen preferred in Russia and now is also applied in US. The liquid may also flow alongrectangular channels machined or formed into a liner fabricated from high-conductivitymaterial (like cooper alloys). The example of such design is the SSME. In US the traditionaldesign have been the thrust chamber and nozzle built from thin rectangular tubes strengthenedby outer bracing (tubular wall). The tubes are directed downwards and upwards the walls ofthe thrust chamber and the nozzle. Since the diameter of the nozzle changes along its axis, theform of the tubes also change (but their cross-section remains constant). The tubes maybifurcate. The hot wall is very thin, so heat exchange with the liquid is very effective. Theexample of this design is the F-1. Efficiency of the regenerative cooling is very high, but there are also some limitations. In 13

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the throat, the diameter and thus the surface of the chamber wall is small, and it is impossibleto pump enough liquid along the limited surface to provide cooling. It is not always possibleto increase the velocity of the flux since it would require raise of pressure that forces thepropellant through the cooling jacket. In this case, film cooling is additionally applied. Themain idea of film cooling is that some fuel in injected through additional injectors into thehotest parts of the thrust chamber right against the wall. The liquid fuel absorbs heat byboiling and evaporating, thus a protective cold boundary film is created and protects the wallfrom contact with the hot gas. This film is spread along the wall by gas moving along thechamber. A variation of this method is transpiration: the coolant gets into the chamber fromthe jacket through a porous chamber wall. Film cooling may be realized also with cold gasfrom turbine directed along the wall of the nozzle to protect it from hot gases from the thrustchamber. The protection of the nozzle extension of the F-1 engine was performed in this way.The injector and combustion stabilityThe propellant is introduced inside the thrust chamber through the injector. The injectorforms a spray of the components to provide their effective mixing an burning. A typicalinjector head consists of a plate with holes for the propellant components organized in aspecial pattern. Some of the injector elements may represent sleeves sticking out from theplate, they may have multiple holes. The injector head may also be divided into sections bypartitions. Stability of the combustion process in the chamber depends highly on the effectiveness ofmixing and thus, on the injector. The size of propellant drops and the parameters of the spraydefine the lifetime of the drops, intensity of their evaporation and the quality of the burningmixture. There is a number of reasons why burning process may become unstable, and thecombustion may become resonant. For example, pressure pulsation may influence the 14

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injection system: raise of pressure inside the chamber slows down the injection rate, and thefollowing drop of pressure (when the excess of the propellant leaves the chamber) leads to anew increase of the injection rate. Self-oscillating process thus establishes with the frequencyfrom tens to hundreds cycles per second. This instability is in strong dependence on thelifetime of propellant drops, i.e. on the delay between the propellant injection and itscombustion. This instability is often eliminated by changing pressure drop on the injector; aninjector pressure drop usually makes ~1/4 of the chamber pressure. Another type of instabilities is high-frequency combustion instability (frequencies >500 cycles per second), it is the most dangerous and is specially pronounced in engineshaving high thrust. These instabilities arise because the time of drops vaporization is notconstant and depends on the pressure near the injector head. The higher is this pressure, moreintensively vaporize the drops, the combustion process accelerates, and a shockwave spreadsalong the chamber, reflects from the opposite wall and returns, raising the pressure evenmore. The period of the oscillations depends on the time in which the shock front returns.These oscillations usually destroy a thin-wall chamber within seconds. This kind ofoscillations is suppressed by changing the chamber length and width, by installing additionalpartitions inside the chamber which divide it into smaller volumes, and by matching theinjector head (number and position of the holes, the sleeves etc.) The designers of the F-1engine, the biggest one-chamber engine ever used, faced the high-frequency instabilities andhad many problems with them. The problems were solved by matching the injector head: itwas divided into sectors by partitions. Vaporization of the components before introducingthem into the chamber also may be applied. A radical way to solve the problem is to replaceone large thrust chamber by several smaller ones: due to smaller dimensions of the singlechambers, the danger of appearance of high-frequency instabilities is smaller. This way waschosen by the designers of RD-170/171: the engine has 4 chambers with the thrust of ~200tons per 15

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chamber. Smaller size of the chamber also permits to decrease the length of the engine. High-frequency instabilities make creation of high-thrust LREs quite problematic and expensive. There is another instability, low-frequency combustion instability (typically 10 – 100cycles per second), not directly related to the injector. It appears due to resonance between thethrust caused by changes of the fuel flow rate, and and proper frequencies of the tanks andstructure of the rocket. The result is so-called pogo oscillations of the rocket. This kind ofoscillations is often eliminated by dampers in propellant lines, like in the Space Shuttle andSaturn V. To damp oscillations, a small amount of helium is introduced into the propellantline to shift the natural frequency of the line and to destroy the resonance.Throttling of the engine. Start and cut-offThrottlingIn general, LREs are throttled by adjustment of the amount of the propellant delivered into thethrust chamber, and this idea may seem to be quite simple. However, in practice it is not sosimple to realize. If the engine is pump-fed, we should take into account the fact that theturbopumb is driven by the propellant, and if the flow rate is decreased, there may be lack ofreaction mass to power the turbine. The pumps are also designed for a certain propellant flow,and significant change of the flow rate may lead to significant drop of efficiency, and theturbopump assembly would work in unbalanced conditions: the power of the turbine may beinsufficient to drive the pumps. Of course, these problems may be overcome technically, butthat would mean impossibility to provide optimum conditions for the turbopump: aseverywhere in technical sciences, an universal unit is usually not so effective through thewhole wide range of working conditions as an specialized unit could be. 16

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When the problems of the turbopump are solved (or in the case of a pressure-fed engine),problems with cooling jacket may arise (if cooling is regenerative). Drop of propellant flowrate may slow down and liquid may begin to vaporize, thus the engine would stall. Theamount of the coolant available also would drop. The temperature in the thrust chamber willremain nearly unchanged if the mixture ratio is intact, and it may become impossible to cooldown the walls. A possible solution is to change mixture ratio and thus to decrease thetemperature in the thrust chamber, but this would lead to decrease of Isp (which would dropanywhere since the pressure in the chamber will drop). Another solution is to use only passivecooling methods (ablation), but that would mean lower temperatures and a shorter lifetime ofthe engine. But the turbopump and cooling are not the only issues to take into account. The mayorproblem is combustion stability. When the propellant flow rate decreases, the injectorpressure drop falls quicker than the pressure inside the chamber, and, as we have seen, theflow rate becomes dependent on small variations of pressure in the chamber. A feedbackbetween the chamber pressure and the propellant supply appears and combustion becomesunstable. To avoid that problem, variable-geometry injectors may be used: the area of theinjector head is decreased when the flow rate drops (this was the case of the Apollo LMDescent Stage). Generally, most of LREs may work with moderate throttling by several percent withoutstalling nor serious loss of efficiency. But deep throttling requires special designs and is adifficult problem to solve. The highest throttling range among the human-rated engines wasthe TRWS for the Descent Stage of the Apollo LM. It could be throttled down to 10%.However, the range of ~65% ÷ 95% was unusable due to stability issues. SSME is throttled in-flight in the range of 65% ÷ 105% (and a little bit wider range is available). RD-180 (a two-chamber version of RD-170) is throttled in the range of ~40% ÷ 100%. 17

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Engine startTo start a LRE, several operations should be performed in a right sequence. First of all, thetanks should be pressurized. This is done with separate gas (like helium stored in specialvessels) or vaporized propellant components. In the last case the vaporized components areavailable only after the engine is started, so temporarily some other gas is used. If cryogeniccomponents are used, specially LH2, the plumbing should be chilled with a small initial flowof the component before the full flow is opened. This is done to prevent boiling of thecryogenic component inside the plumbing. If turbopump is applied, the turbines should be started to begin delivery of the propellantinto the chamber. This may be performed by burning main components in the gas generator:the initial amount of propellant is directed into the gas generator to gasify. Turbines may alsobe started by initial flow of a separate gas like helium stored in a starting vessels. Whenturbines are able to pump components, the flow of the propellant into the chamber is initiatedby opening main vents. If propellants are not hypergolic, they should be ignited. For a single-burn engines, likethat of launch vehicles, chemical ignitors are often used : a small amount of hypergolicpropellant is introduced into the gas generator and the thrust chamber before the mainpropellants, and they ignite the propellant mixture (often this hypergolic component self-ignites when mixed with the main oxidizer). Ignition with a pyrotechnic charge or even atorch introduced into the thrust chamber also may be applied. For multiple-burn engineselectric ignition (with spark) is often used. To avoid hard start (when the pressure in the thrust chamber rises too quickly, damagingthe chamber), the start of a large engine should be performed carefully. Sometimes the start is 18

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performed in two stages: at first the engine is started at a fraction of the full thrust, and whenthe thrust is raised to the nominal value. The engine may also be started at a mixture ratiodifferent from the nominal, and then the ratio is set to the nominal.Engine cut-offLarge engines cannot be cut-off by only closing main propellant valves. Since turbinescontinue rotating for some time, the pressure in the pump tract will not drop immediately. Toavoid raise of pressure at the main valves, the flow may be redirected to a low pressure line. Ifthe engine is pump-feed, at first the propellant flow to the gas generator is decreased, and theturbines rotation slows down. Large engines are often cut-off in two stages to avoid rapidtransient processes that might damage the engine and the rocket. Pressure-fed engines may becut-off by propellant depletion: the flow of one of the components stops, and the engine shutsdown itself. 19

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Physics of a rocket engine and nozzleIn the previous lecture we have seen that the thrust T of a rocket T = qVmean + S a ( pa − p )engine is defined by the formulaHere q = dm/dt – mass flow rate, S – area of the nozzle, pa – static pressure of the exhaustgases, p – ambient air pressure. If pa > p, the nozzle works with underexpansion, if pa < p,the nozzle works with overexpansion, pa = p is the case of ideal expansion. Increase of theexpansion ratio ε (the ratio of the pressure inside the chamber pch and the pressure at thenozzle exit pa, ε = pch/pa) leads to an increase of the first summand and to a decrease of thesecond summand. It may be proved that the case of ideal expansion provides the highestthrust possible for the corresponding external pressure. As the rocket ascends, the ambientpressure changes, so it is impossible to provide optimal conditions for all heights. A nozzlethat has overexpansion at lower heights (after the lift-off) will work in underexpansionconditions after the liftoff. In vacuum, every nozzle works in underexpansion conditions,since it is impossible to provide zero pressure at the nozzle exit (for that, the nozzle expansionshould have infinite length and with). But larger is the nozzle, less is the underexpansion.However, it makes no sense to increase very much the size of the nozzle, since the gain ofefficiency would be cancelled by grow of the nozzle size and weight. For the first stage, theexpansion ratio of the nozzle cannot be very large, since too high overexpansion would leadto separation of the jet of the exhaust gases from the nozzle. That may lead to catastrophicdrop of efficiency and even damaging the nozzle, if measures are not taken. Generally, thenozzle of the first stages work in overexpansion conditions in the lower atmosphere. Nozzlesof upper stages work at significantly lower ambient pressures, so they 20 much larger haveexpansion ratios than the

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nozzles of the first stages. The expansion ratio of a nozzle is a compromise between thrustefficiency at different heights, weight and size. In the recent time an original design havebeen implemented in several engines: the nozzle extension changes inflight to improveefficiency while the rocket ascends from the denser atmosphere to more rarefied layers. Thissolution is applied on the second stage of the Delta IV rocket (LOX/LH2 engine RL-10B-2). The shape of the rocket nozzle is of de Laval type. It is convergent-divergent, having athroat. In the convergent section of the nozzle the gas flow is subsonic, it velocity increasestowards the throttle and achieves the sound speed at the throttle. In the divergent section thegas jet expands and achieves supersonic speeds. The gas escapes in slightly different directions in different parts of the nozzle. This factmay be taken into account with the coefficient ϕ that relates the Vmean 1 + cos αescape may velocity to the actual escape velocity Va and the half-angle of Vmean = Va 2conicity of the nozzle α: 2λ RTch   λ −1 From gas dynamics, the gas escape velocity may be found as Va = k 1 − ε λ R – gas constant, ε – expansion ratio, T – temperature inside the λ −1 µ     chthrust chamber, λ – adiabatic index, µ – molar weight of the exhaust gases, k – imperfectioncoefficient that takes into account incomplete combustion, film cooling etc. It is seen that higher escape velocity are provided with higher temperatures in the chamberand lower molar weights of the exhaust. The dependence of the expansion ratio has a 21

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maximum at ε = 0, i.e. the pressure inside the thrust chamber should be possibly high and theouter pressure should be possibly low. The term in the brackets is η – expansion efficiency, Vc = (λRTch /µ )1/2 V = kV 2 a c ηis the sound speed. So formula the escape velocity may be given a form λ −1 For typical propellants λ = 1.25, and the escape velocity is ~2.8 Vc (if k ~ 1 and η ~ 1). 22

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Injector headsInjector head of the HM-7engine (LOX/LH2, Arianeupper stages). Sleeves for oneof the components are insidethe holes for the anothercomponent (By source) Injector head of the RD-170/171 engine (LOX/kerosene, Energia/Zenit first stage). Sleeves for one of the components form partitions (By source)